text:
In wing design, how is factor of safety related to G-limits?
What does factor of safety mean with regard to the maximum G force an aircraft is rated for? Does a rating of 3Gs mean the same as a safety factor of 3? Or is there some extra margin?

what context are you talking about?

@Federico trying to figure out loading for spar

Structural Factor of Safety is the margin between the normal limit load and the ultimate load the structure has to take.  It is normally 1.5 times the yield load or limit load of the structure.
The structure is designed to withstand the Limit Load that is applicable to the category, in the case of Transport airplanes 2.5G, without permanent deformation.  The Factor of Safety is applied on top of this value as a minimum, below which the structure may permanently bend, but not break completely.
So an airliner has to be good for 2.5G without permanently bending the wings, with a Safety Factor value of 3.75G before the wings break off, and you have to design your structure taking into account of the elasticity and ductility of the material, to achieve both those requirements.
The Limit Load is the maximum load the airplane is expected to encounter out in the world doing its thing.  The 1.5 Safety Factor been been in use since the 1930s and is a fudge factor to account for all sorts of variables, like assembly defects, calculation errors, aging damage etc.
So to answer your question, although you said 3G, we'll say the value you're interested in is the value for light planes in the Normal Category, which is 3.8G, which would be the design limit load that it's expected to be able to take with no permanent damage.  The Safety Factor, which by industry convention is 1.5, is added on top of that value to get 5.7G and is expressed as design ultimate load for Normal Category. So you have to make sure your wing can take 3.8G and still return to its original shape, and 5.7 Gs before breaking off.
You try to get as close to these limits as you can in your design, because every pound of weight above the minimum requirement is wasted ballast as far as the airplane operator is concerned. However, designers may add additional structure on top of that calculated to meet the regulatory minimum just to be safe (especially if the materials are new), or to improve fatigue life. This is why wing structures typically get tested to failure after their fatigue test program is completed, to see what will happen and to prove they can meet ultimate load requirements, and usually fail well above the design ultimate limit.
You might find this report from the 1950s helpful.


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If sweep delays compressibility, why does 747 cruise slower than A380?
B747 has a sweep angle of 37 deg and A380 has a sweep angle of 33 deg. If the main reason for sweep is to delay compressibility effects and increase operating Mach number, why does B747 cruise at a lower speed than A380?

This appears to only be true for the old 747

From Wikipedia, the cruise speeds of the both airplanes are fairly similar:

A380, nominal wing sweep angle 33.5 degrees, cruise speed Mach 0.85

Boeing 747: nominal wing sweep angle 37.5 degrees, cruise speed Mach 0.84 .. 0.9, depending on the model.

So those numbers would contradict the claim in your question, but align with your expectations.
There are several other things to consider. The nominal cruise speed of these planes is more dictated by efficiency, required power for a given speed, and fuel prices. The 747 is an older design, when fuel was cheaper, the general tendancy is that modern airplanes reduced the cruising speed slightly in favor of higher efficiencies.
considering the sweep, raw numbers:

cos(33.5 deg)= 0.83
cos(37.5 deg)= 0.79

So the difference of approx 4% here seems to reflect the difference in cruising mach numbers: the higher sweep of the Boeing 747 reduces apparent Mach number roughly by the same amount that it is flying faster.
Then, for both airplanes, and certainly for the A380, the wing aspect ratio is quite low. So there's a point to challenge how much the sweep actually angles differ, from an aerodynamic point of view. It may be very well that effective angles depend a lot more on local airfoil shapes, low aspect ratios, and other 3d effects than they depend on the nominal sweep angle.
All that said, the swept wing has more advantages than just reducing the apparent Mach number. There is

enhanced aerodynamic stability around the yaw axis (this may even help in the event of asymmetric thrust due to an engine failure)
enhanced mechanical stability of the wing itself under aerodynamic loads
trim properties in transsonic aero regimes

The final design is then a result of many different considerations.
Eg, IIRC, the wing span of the A380 had a limitation so it would "fit into standard airports", and there were likely some trade-offs in aerodynamics for that reason.

The A380 had to hit the Mach 0.85 sweet spot in order to show up on the first page of booking systems, see https://aviation.stackexchange.com/questions/11936/are-we-at-peak-speed-efficiency-for-jet-airliners-at-mach-0-85/11948?r=SearchResults&s=5%7C5.2899#11948.

It's important to distinguish between three different Mach number definitions:

Design cruise Mach I believe is 0.855 for the 747-8 and 0.85 for the 380.
Maximum Operating (or Mmo) is 0.92 and 0.89, respectively.
Demonstrated Maximum Diving (or Mdf) is 0.97 and 0.96, respectively.

Multiple factors effect drag from compressibility effects, especially the overall aircraft configuration area distribution along the longitudinal axis (ref. wave drag).

The question implies:

that the maximum flying speed is mainly affected by the compressibility drag and;
that compressibility drag is mainly related to sweep angle;

but this is not the whole story.

why does B747 cruise at a lower speed than A380?

Maximum flying speed depends on the equilibrium between thrust supplied by the engines and drag generated by the aircraft.
Thrust obviously depends on the technological development of the engines (which was for sure different between the first version of the B747 and the first version of the A380) and, among other factors, on temperature, density, humidity, height, ...
Drag depends as well on the technological development of (the surface of) the aircraft: its shape, how much surface can maintain a laminar boundary layer, surface finishing, eventual use of area-rule, type of airfoil used for the wing, ... Plus (induced) drag depends on lift as well.
So, maximum flying speed is affected by a lot of parameters and not only by compressibility. And upon this, also structural and/or aeroelastic (aka flutter) limitations can restrict maximum speed.

the main reason for sweep is to delay compressibility effects.

Compressibility drag is not influenced by the sweep angle alone.
If for example the area-rule concept has been used in the design phase then the geometry of the whole aircraft is important (area-rule is a particular smooth variation of the sectional area of the aircraft which minimises compressibility drag at transonic speed. Sectional area is the area that you get when slicing the aircraft longitudinally).
Another way to reduce compressibility drag is the use of supercritical airfoils, which have been for sure employed on the A380, not sure on the B747.

So, sweep angle is not the only parameter influencing the value of maximum speed or compressibility drag and its value is actually chosen as a compromise of many other (contrasting) requirements.

Bonus material about sweep angle: the point on the airfoil where the highest speeds are reached is around where its thickness is the highest. For a jetliner this point is important because it is the point where speed can reach transonic or supersonic speed with a relevant increase in drag. For this reason the real sweep angle to be considered for aerodynamic considerations is not the one at the leading edge rather the one passing through the highest thicknesses of the wing.

source: Stack Exchange

license: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/4.0/

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Which manuals should I study before flying to France, Portugal and Switzerland?
I am from Brazil. My crew and I will fly to France, Portugal and Switzerland. I am accustomed to flying to the USA, but it will be my first time to Europe, and I am studying the differences. Which documents and manuals do you recommend that I study and read? (Note: I want names not resource location)

Doesn't your company have materials to prep you for entry into these countries? Maybe I'm not getting it clearly, but are you saying that you and your "crew" are the airline crew, or some other kind of crew?

We are corporate pilots.

We do have materials, and they will pay a course too, but we'd like to read more about procedures, contingencies and etc.

I am looking for the "International Route Manual" in which there are details about some aiports there.

There is a lot to learn, especially about the crossing but also about the differences in the way that they do things.  I would highly recommend taking an International Procedures course. These are offered by the major training centers, as well as various online / home study courses.

All these country are member of EASA (equivalent to FAA in europe) and apply it's rules, i'm not very concerned about Air Ops but I think you can find more information here : https://easa.europa.eu/easa-and-you/air-operations

For common procedures and or VFR rules you want to check the respective countries' AIP:

Switzerland   http://www.skyguide.ch/en/services/aim-services/
the other countries you mentioned can be found from this link collection http://www.eddh.de/equipment/ais.html

you may want to check the website of Eurocontrol here you can find IFR standard arrivals and departures

No warranty! Please prepare your flight carefully.

all_licenses: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/3.0/

all_licenses: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/4.0/


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I need to know how to calculate drift with a Jeppessen CR-3. Can someone help?
In my exams we are not allowed to use E-6B, just a CR-3 jeppessen flight computer and I'm having serious issues calculating W/V from other parameters like TAS/GS and HDG/TRACK.
Example:
Q : Given:
True HDG = 035°, TAS = 245 kt, Track (T) = 046°, GS = 220 kt.
Calculate the W/V?
A: 340/50kt
I need to know how to calculate Track HDG and Drift Can someone help ?
HDG-001degree TR-? Drift-05degree
What should be the Track? and what will be the Steps to solve !


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How does gas turbine efficiency compare with supercharged piston engines?
If an aircraft mounted the exact same propeller on a supercharged piston engine and a turbine, and flew under identical conditions at the same pitch and rpm, how much more fuel would the turbine use (possibly including weight savings benefits for extra credit).
I am picturing a two engined aircraft with the two types of engines flying side by side at around 350 knots.

possible duplicate?  https://aviation.stackexchange.com/questions/50768/which-jet-engines-have-the-highest-thermal-efficiency

The most efficient IC engines are large Diesels. At the extreme end are ship engines with better than 50% thermal efficiency resulting in a specific fuel consumption of only 0.260 lbs/hp/hour or 158 g/kW-h. But even supercharged truck diesels achieve above 40% thermal efficiency at high load (this NHTSA study gives 42%).
Aerodiesels have achieved 220 g/kW-h already with the Jumo 204 and 205 of the early 1930s. Even the modern Thielert diesels (now sold by Continental) are hardly better, claiming 214 g/kW-h. Also the Napier Nomad, a super- and turbocharged aero diesel with utmost efficiency as its design goal just reached 219 g/kW-h.
Gasoline engines start at about 240 g/kW-h; this value is achieved by the Lycoming IO-390 with fuel injection. Without fuel injection, specific consumption rises to 260 - 280 g/kW-h which is typical for a Lycoming O-360 at 65% power. Note that the Jumo 213, one of the more efficient WW II-era piston engines, already achieved 260 g/kW-h even with 87 octane fuel and a compression ratio of only 6.93:1 at its most favourable operating point. Advanced Innovative Engineering, who have taken over the Norton-Wankel engine, claim 310-350 g/kW-h for their 650CS with 120  PS.
Comparing that with turboprops needs some conversion of thrust into power. This is only valid for a specific flight speed. If you do that at cruise speed, the large turboprops Progress D27 and Europrop TP400 claim a consumption of around 240 g/kW-h. Smaller turboprops rarely achieve below 300 g/kW-h.
To save you from the trouble of looking up and converting the data in the last link, here is a selected list:

Allison 250 $\;\;\;\;\;\;\;$: 370 g/kW-h. This is a typical small helicopter engine.
Garrett TPE331$\;\;$ : 310 g/kW-h. This is used on small turboprops like the Do-228 or the Merlin III.
PWC 126A $\;\;\;\;\;\;\;$: 280 g/kW-h. Getting larger - BAe ATP.
Rolls-Royce Tyne : 237 g/kW-h. This has long been the largest turboprop in the West and used on aircraft like the Canadair 400 / CL-44.

Please note that those turboprops feed on kerosene while the piston engines need gasoline. But by basing the comparison on a per-mass basis, it is valid because the energy densities of both are almost identical. Very large turboprops are as efficient as gasoline piston engines, but diesels still have a small advantage.
Now for turbofans. Here we have thrust which needs to be converted into power first by multiplying it with flight speed. It would be nonsensical to compare the static case – here by definition turbofans do not produce power. For the nitpickers: Yes, I need to look at the gas speeds ahead and behind the engine, but still, this makes for a poor comparison: Most static values are from test stands with all accessories removed and no losses for engine mounts and fairings. I will instead use the figures in cruise given in this answer, using a fuel burn of $b_f$ = 18 g/kNs and a speed of Mach 0.78, which equates to a flight speed of 262 m/s in 11.000 m altitude. Multiply by 3600 for a per-hour value and divide by 262 (the N are in the denominator!) and you arrive at 247 g/kW-h. So again, very similar to good gasoline piston engines but not as good as diesels.
But again this comparison has to be taken with the proverbial grain of salt. Now we need to have a closer look at speed. Thrust-specific consumption goes up with speed and roughly doubles between the static case and cruise speed for a modern turbofan. The GE-90 achieves 8 g/kN-s static and 15 g/kN-s at Mach 0.8 – that would be just 209 g/kW-h and be on a par with the best diesels. For comparison: The installed figures for modern military engines in supersonic aircraft are 20 g/kN-s. And regarding the fuel-hungry turbojets: The old Jumo 004 achieved 39 g/kN-s – just twice as much with a compression ratio of only 3.3:1. Real fuel guzzlers were the Argus 014 of the V-1 with 107 g/kN-s in cruise.
While turbine engines gain in efficiency with altitude due to colder intake air, the diagram below comparing the Jumo 213 A with the J version (source) shows an increase in power specific fuel consumption with altitude. Note that flight speed will also rise with altitude and is not given, so I suspect this is more due to higher speed than higher altitude. Again, these are real-world data from flight test with the engine installed in a FW-190D (source). Going from sea level to 10 km which roughly doubles true air speed raises the specific consumption by 20%.

Comparison chart between Jumo 213 A and J. Flight altitude is given in [km] along the x axis and specific consumption along the right y-axis. Multiply by 1.34 for g/kW-h. The lower set of lines are for partial load operation between 2100 and 2700 RPM (A version) rsp. 3000 RPM (J version) while the upper set of consumption lines are for maximum power operation at 3000 RPM (A version) rsp. 3400 to 3700 RPM (J version), partially with water-methanol injection.

Kampf absolutely outstanding once again.  The Tyne generates around 6100 hp.  I am seeing a "reverse Brabazon" with one, larger, more efficient jet turbine powering 2 props.  Can we go a little higher (even with a fan)?

@PeterKämpf I'm guessing he's saying it means "gallons".

A kilowatt is a really strong light bulb.  1 gallon per hour for that?

@RobertDiGiovanni: Gallon is so obviously wrong it hurts.

@Roger please clean this up

Sorry for the confusion.

How does turbine efficiency compare with internal combustion engines if all the turbine power is converted to mechanical energy?

When looking at conversion from chemical energy into mechanical energy: very favourably. The early turbojets had low thrust efficiency, they could not convert their gas generator power into thrust in an efficient way. When comparing energy efficiencies we start with the chemical energy in the fuel, and won't be working backwards from available thrust.
If all gas generator power is converted to mechanical energy we're talking about turboshafts. The best way to compare the fuel efficiency of the engine only and not get engaged in a discussion about thrust conversion, is a listing of engine brake horsepower, which directly measures the shaft torque1 applied to a propeller, fan, set of truck wheels, ships propeller etc.

As can be seen, a large gas turbine engine like the GE LM6000 (a converted aeroplane turbofan engine) is among the most frugal engines around, with a brake specific fuel consumption of 0.329 lbs/(hp * h) = 200 g/kWh = 42% efficiency.
Diesels can get efficiencies of over 50%, mainly the large, low RPM 2-stroke diesels generating a huge amount of torque. The Wärtsilä-Sulzer RTA96-C runs at 22-120 RPM, which would be problematic for automobile and aircraft propulsion.
The highest efficiency of all is the combined cycle with 62.2%, a gas turbine combined with a steam turbine to utilise the exhaust energy.

The low specific fuel consumption is for large gas turbine engines, they don’t scale down well because of the boundary layer effects: a smaller engine has relatively a larger circumference. Also, they only operate efficiently at full power, piston engines have the advantage at lower rpm percentage.
So the smaller the engine, the more advantageous are the circumstances for the piston engine: they scale down much more favourably than the gas turbines. But given enough volume, gas turbines are not inherently wasteful with fuel at all.

If an aircraft mounted the exact same propeller on a supercharged piston engine and a turbine, and flew under identical conditions at the same pitch and rpm, how much more fuel would the turbine use

We're taking about aero engines now, and should leave unsuitable 2-stroke low RPM ship diesels out of the comparison. Also, since the propellers and flight conditions are identical, we can leave the whole thrust conversion mechanism out of the equation. As mentioned earlier, size matters when talking about gas turbine efficiency. Let's take 2 sizes:

Largest turboprop.
The absolute largest turboprop engine was the Kuznetsov NK-12, developed just after WWII. A more modern large turboprop is the Europrop TP400, only slightly smaller, used for the A400, 2 with a maximum power output of 11,000 hp = 8,203 kW.

Cruise shaft power specific fuel consumption: 0.167 kg/kWh (0.275 lb/hp/h) = 51.5% efficiency.

Cruise propulsive power specific fuel consumption: 0.213 kg/kWh (0.350 lb/hp/h)
But it may be seen that large gas turbine engines are very fuel efficient, and are not inferior to piston engines in frugality.
I could not find fuel consumption data on a comparable piston engine that delivers 11,000 hp. The only reference is to a drag racing engine, which is not very concerned with fuel economy.

Largest piston powered prop
The largest piston engine ever produced was the Lycoming XR-7755, producing 5,000 hp at a fuel economy of .38-.41lb/hp-hour, 231-249 g/kWh or 35.5-32.5% efficiency. And that with 1947 technology.
Also from the 1940s: the Allison T40 turboprop with 5,100 hp and specific fuel consumption at 0.63 lb/(she*h) = 383 g/kWh = 22% efficiency. Low compression ratio, no FADEC.
The Q400 has two PW150A turboprops of 5,000 hp each. Could not find the SFC of this modern gas turbine engine. The Rolls Royce Tyne is from the 1950s and does not have the latest technology.

It is hard to find equivalent piston and gas turbine engines for a fair comparison:

Time of development. Avgas pistons were highly developed after WWII, afterwards development of large engines stopped, now only used in smaller aeroplanes. Turboprops now have 3-axes, very high compression ratio and inlet temperatures, FADEC.
Engine rated horsepower. Turbines don't scale down well, and the comparison HP rating will have a large influence on the outcome.

1: The gross power delivered by the engine is the torque Q at the shaft times the angular speed $\Omega$. This is the pure engine power, the Shaft(!) Horse Power of piston engine specifications. To measure the gross shaft power, connect an eddy current damper to the output shaft and measure the amps required to run at constant $\Omega$
2: By Matti Blume - Own work, CC BY-SA 4.0, https://commons.wikimedia.org/w/index.php?curid=68573331

The other big disadvantage of piston engines is the need for different levels of supercharger boost at different altitudes. If you look at efficiency in terms of miles per gallon, the turbine engine's ability to fly high gives it a big advantage.

@Robin Bennett the B36 went higher than many people know (actually may have had a TAS close to the B52).  "Turbocharging" really is the same for jets and pistons, it's the combustion and thrust efficiency that differs. (Still trying to get them to figure AVERAGED piston force as "thrust" vs the "drag" of the prop airfoil.  But weight vs thrust, maintenance, and max power, turbines win hands down.  I want one running my four props (maybe electrically).

@RobertDiGiovanni The average piston force is what is found back at the shaft, as torque. The pulses integrated over time result in a force, multiplied by the crankshaft radius result in a torque. The shaft torque in a turboshaft is the result of the continuous sideway force of the gas stream on the turbine blades, easier on the intuition.

And lots more power too.  But the piston force is from the (more efficient) internal combustion expansion.  Just too heavy and can't get enough of it.  Interesting to compare heavier cannon to recoiless "rifle".  Cannon are more efficient, but harder to carry.  Amazed at how good the larger turbines are though.  Food for thought, the piston force is straight, the turbine oblique, but multiple rows of turbine blades seems to help.

Now we can add distance travelled to g/kW-hr to get ton-miles/gallon like efficiency, which is the whole package including optimal speed and altitude.  The 747 is the best I've seen yet at 80 ton-miles/gallon.  Need to covert that to g-meters/kW-hr.

Well the turbine gases expand as well, following the same Carnot cycle as stated in @ic_fl2 answer. You're right though in that the piston has no pressure losses during expansion - on the other hand, it does have friction losses. All in all the Carnot cycle parameters dominate: higher compression ratios for diesel pistons and turbines.

@RobertDiGiovanni Need to covert that to g-meters/kW-hr That is where a whole matrix of parameters is introduced which have to do with flow propulsion, not with engine efficiency.

Yea, friction loss a great point for pistons, 4 cycle do better than 2.  Diesel seems most like jet in that combustion is done with excess air. Large pistons have issues with heat removal, jets only get better (validation of history).

Very poorly actually.  What saves turboprops and turboshafts is power to weight, smoothness and reliability.  If you want just maximum MPG and don't have to go trans sonic, recip wins hands down.
Piston engine Specific Fuel Consumption is roughly .45 lbs/hp/hr for a normally aspirated carbureted engine (that figure comes from my own Lycoming engine's power chart - I can't find an on-line source), going down from there with fuel injection, and turbo/super charging. Diesels are in the low .3s.
The most efficient recips were the Wright Turbo Compound radials that had both supercharging and direct horsepower extraction from the exhaust (about 300 hp was recovered from the exhaust of the R3350 by the two power recovery turbines) that were down in the high .3s to .4.
Turboprops?  Worse than two stroke gasoline engines. Somewhere around .6 lb/hp/hr or worse (the PT-6 is .67), maybe in the high 5s on some of the latest.
This is one reason why you don't see that many turbine conversions on airplanes like the DC-3.  Aside from the cost of the conversion, it's just way more economical to run, fuel wise, on gas.

Have you got any references for those numbers?

@Koyovis I put a couple links to the Wiki page for the 3350 and PT6 which lists their numbers.

Turbine engines have similar fuel efficiency to diesel engines - if they are large enough. Check out for instance the GE LM6000 in this Wiki page.

@Koyovis Comparing apples with oranges again? The biggest diesels easily surpass 50% thermal efficiency. But they are heavier than turbines, which makes turbines attractive to short-range aircraft: Helicopters!

@PeterKämpf  The 2-stroke diesels are listed as the most efficient engines, large turbines cannot match that. To claim that only power-to-weight makes the turbines suitable for aircraft would be missing the point though.

As can be seen from [the Wiki page], which lists the GE LM6000 as 0.329 lbs/hp/hr.

@Koyovis: I strongly suspect that those Wikipedia values are for the naked engine on a test stand. With accessories removed as much as possible. Real-world values are quite a bit higher.

@PeterKämpf So comparing apples with apples, and listing purely the fuel efficiency of the core engine. Which is what the OP asked.

I figured the OP was asking about a comparison between comparable aircraft engines on airplanes with comparable performance, basically recip vs turboprop.  Recip wins in that case.

@JohnK I understand, I was looking at the title.

@JohnK don't you want to edit the post?

Very well as the Carnot efficiency is what drives the theoretical limit, and that is driven by the temperature of the combustion. 1 - T_h/T_c (in Kelvin) Turbines have very hot combustion chambers (diesels too, gasoline less so (due to knocking)).
So a turbine operating in cold air will have great performance.
In electricity generation Natural Gas turbines combined cycle (CCGT) GTs that are coupled with a steam turbine achieve 60%+ efficiency. Turbines only (OCGT) have lower efficiencies (34%) but that is also because they are optimised for faster response to power prices.

Your last paragraph seems a bit of a tangent: are there many natural gas turbines in aviation? Could you [edit] your answer to make the numbers a bit more specific to aircraft turbines?

Edited to clarify. The initial question is about all turbine power to shaft, which is the case in the electricity generation, but not in aviation as you can use the energy in the exhaust for propulsion directly. I guess you could have a closed cycle steam turbine and pug the whole thing into electric motors... Large ships diesels as per the other answers are not used for air plane propulsion either, but in the search of efficiency there are some crazy designs.

A large Bergen diesel will use 175 g/kWh at 85% of max load. At 600 kW per cylinder so a v16 is 10 MW or 13,000 shp. It will also weigh 150 tonnes, vs a TP400 gas turbine that weighs 2 tonnes and generates 11,000 shp at 210 g/kWh
https://www.rolls-royce.com/~/media/Files/R/Rolls-Royce/documents/marine-product-finder/diesel-and-gas-engines-brochure-1216.pdf
https://en.wikipedia.org/wiki/Europrop_TP400

I don't see how these figures provide an answer to the stated question.

@RalphJ why? At worst, this looks like a promising answer

Actually, this is an excellent answer!  Within the context of the original question, this answer provides some perspective.


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What are the takeoff and landing distances of motor gliders?
I'm doing some estimations of weight and thrust loading for a college project, and in the formula below the landing distance is not a variable, but should be estimated using similar aircraft and usual runways. (There is also a formula for landing with the same problem)

The similar aircrafts are 18-21 meter self-launching motor gliders like DG-505MB, DG-1001M or ASG 32 Mi (but with absolutely no data I have now, information about any other gliders would be appreciated as well).
The only thing I've found was that European Regulations CS-22 allow motor gliders a maximum takeoff distance of 500 metres.
Source; T.C.Cork, Design of Aircraft

Check the flight manuals, there should be plenty online (e.g. https://www.williamssoaring.com/fleet/ASG32_manual.pdf -  the ASG 32 Mi needs 470 m to reach 15m taking off a hard runway)

Thanks! I just haven't thought of looking somewhere else than manufacturer's website. There seems to be no information about the landing distances though.

The landing distance of gliders is typically smaller than take off distance and therefore not really of interest. I've looked and some motorgliders like the G109 have it, but I couldn't find powered sailplanes with that infomation.

You may wish to obtain POH data from the 18-21 meter motor gliders you have listed and compare it to your calculated results. Also look into information such as wing loading, aspect ratio, takeoff speed (or stall speed), and airfoil type in addition to power source to include in your report.
Motor gliders, true to their name, should lift off well before 500 meters.  Landing distance would be similar and also available in POH.

Thank you! I thought that this was some secret data but it's just the fact that manufacturers don't post their manuals publicly.
I found plenty of takeoff distance data, but there seems to be no information about landing distances. Would you think it's safe to assume equal landing and takeoff distances?

I would not assume equal, comparable might be a better term, as takeoff can be more precisely measured.  Generally, landings include how far down the runway, ground effect, flaps, braking etc.  Since you are working with weight and thrust equation, why not focus on takeoff and use those distances for starters.  Curious, what exactly are you trying to calculate?


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Why are no recent small aircraft designed to be "characteristically incapable of spinning" as the Ercoupe was?
I'm a student pilot at a flight school that primarily flies RV-12s as a trainer. We had one day that through a booking error meant that all the RV-12s were booked so we had to use their Ercoupe 415-C instead.
I was surprised at how easy the plane was to fly and asked my CFI about the "characteristically incapable of spinning" placard on the front panel. He explained that the rudder pedals weren't necessary because the plane maintains coordinated flight by connecting the ailerons to the rear rudder.
I think this is an amazing innovation for the Ercoupe and I'm surprised it's not a design that I can find on any modern small aircraft designs.
Maintaining coordinated flight with rudder pedals isn't a complicated task for a trained pilot but it is another task that the pilot needs to perform. It seems like removing flight tasks from the pilot would make for a safer aircraft so why do manufacturers not include designs similar to Ercoupe's models?
Is it a weight/cost/drag issue, or are spins not as huge of a killer as they were in the 40s-60s?

Because taking away the rudder pedals harms pilots' fragile egos.  And as long as you have rudder pedals, you might as well have a spin-prone aircraft.  That's good for pilots' egos as well.

Depending on speed, you need different ratios of aileron and rudder for coordinated flight. If you always fly at the same speed, coupling both is fine. However, you will not be able to sideslip into a tight field. And slipping is fun!

With regard to the portion of your question about modern small airplanes, Cirrus aircraft feature a rudder-aileron half-interconnect. Applying rudder will move the ailerons, but applying aileron will not move the rudder.  And yes, the interconnect can be easily overridden with muscle for crosswind landings.

Because the amount of rudder necessary to maintain "coordinated" flight cannot be determined by any simple mechanical system that relies solely on aileron input. It is dependent on AOA, (NOT SPEED, @Peter), pitch rate, flight path angle rate, yaw angle, etc. etc. Even in very sophisticated modern aircraft, the computers that are used to do this can screw it up in ways that a pilot would not. Also, there are times when you want (NEED) to be able to fly uncoordinated.

Another interesting approach (which also had it's issues), was used in early F-15s. The effect of pushing the ailerons to one side or the other was different based on the fore/aft position of the stick. When the stick was fully forward, (implying low AOA), lateral movement deflected the ailerons on the wings. When it was far aft, (implying high AOA), lateral  movement caused little or no aileron deflection and almost all differential stabilator. Median fore/aft stick positions would cause a mixture of both.

@CharlesBretana It's neither speed nor AoA but lift coefficient. But why make things more complicated than necessary? Please note this is a student pilot, not an academic. In the end, speed is easiest to observe and the dependency should be obvious.

@Peter, Although everything in Aerodynamics is complex, such that All parameters, properties, characteristics, whatever have multiple interdependent relationships with each other, the consistent relationship between AOA and the Lift coefficient (represented by the Lift curve), is probably one of the most consistent, and reliable relationships (if not the Most consistent/reliable one), of all.

(con't) ... and also, although AOA and the lift coefficient  are tightly bound to one another, no such connection exists between AOA/Lift Coefficient and Speed.  A pilot can establish any Angle of attack within a wide range at any airspeed. These two parameters are not connected in any significant way, (unless you posit an autopilot
or pilot attempting to maintain an aircraft altitude or attitude).  Indeed, the failure of the FAA and the aviation community to focus student understanding on the importance of AOA (in contrast to airspeed), is, imho, a long-standing tragedy.

And finally, your comment that "speed is easiest to observe" misses the mark. In the arena we are talking about, what the pilot observes is not as important as what he/she Feels. And AOA is much easier to feel (sense), and can (could) easily be taught to students. The feel of the stick or yoke in your hands, and its position in your lap is the most obvious. The picture in the windscreen, as it changes, recognizing the point where the aircraft is headed, and where that point lies in the windscreen, is another indication.

The coupling between the ailerons and rudder are designed such that they avoid uncoordinated turns as you described.  However, this occurred at the expense of control in other flight conditions.
In fact, there are situations where you need to deflect the rudder without roll input, you need to deflect the rudder more than the preset coupling, or in a manner that is opposite to the preset configuration.
For example, if you turn right, the system will add right rudder to keep the airplane coordinated. But if you want to do a forward slip, you may perform a left hand roll, but with rudder all the way to the right. The system would be incapable of performing this maneuver, meaning you can't fly in high winds with the system installed. This is fine if you live in an imaginary place where the winds are always favourable.
As a result, a common modification on the Ercoupe is to disconnect the coupling between the rudder and ailerons...
Edit: I also wanted to add that spins are recoverable, within reason, by a trained pilot. For some planes, you are simply forbidden to intentionally enter a spin as a remedy to its inability to recover from a spin!

Very interesting! It makes sense that you couldn't do a slip landing in an Ercoupe. That is certainly a downside.

Also, one of the things most student pilots learn is how to avoid a stall or spin, unless you're practicing them.

Sounds like you want pedals that offset/trim what the automatic system would do.

It also adds more things to go wrong. The coupling jams and suddenly you've got problems with both the rudder and ailerons.

"Inability to recover from a spin"? I was thinking that the FAA regulations require at least all civilian airplanes to be able to recover from a spin using the "standard" method?

Thanks for the great answer on this. I had expected the double vertical stabilizer to have something to do with the stabilization. Is that a part of this at all? What's the benefit of having two vertical stabilizer over one?

Hey @Burke9077, having a twin tail can have bring benefits when it comes to aerodynamics and redundancy for a given aircraft, and is usually done to get a reduced footprint (height) for storage. I am not sure what the design priority was for the Ercoupe, perhaps that is another question to ask! Take a look at this thread: https://aviation.stackexchange.com/questions/35871/why-the-use-of-twin-tails

I have been told that an aircraft's propensity to enter a spin is in fact directly related to the capability to recover from one. i.e., that the harder it is to get a specific aircraft type into a spin, the harder it is to recover from it.

Rudder pedals are used for more than keeping turns coordinated. Forward slips for crosswind landings, slips to increase drag during descents, and reducing the tendency to weather-vane during cross-wind taxiing are examples.

The ercoupe was an advanced design for its time, but it represented a bunch of compromises and the inability to do a forward slip was one of them. The plane's designer made up for this by choosing the airfoil so the plane would descend steeply enough with power off so the pilot wouldn't have to bleed off altitude if he was high on final, and by making the landing gear stout enough to manage landing while severely crabbed.
Another compromise was the omission of flaps; this was to simplify both flight operation and maintenance but required an airfoil that yielded a flap-like descent with power off (as mentioned above) which is why the ercoupe glides like a brick.

It doesn't glide any worse than its contemporaries.  The procedure for losing altitude on final was to S turn.

handbook value for 65HP Piper J-3 is 10:1, 6:1 for the ercoupe...

Yeah but lots of contemporaries, biplanes, the short wing Pipers, also glide like bricks and you still have to side slip them from time to time, and if you had interconnected controls, you'd be left with S turns.  6:1 or 10:1 L/D is neither here nor there.

The usual way to make a plane 'spin proof' is to ensure the centre of the wing (near the fuselage) stalls before the tips, and to move the Centre of Gravity far enough forward that the elevator doesn't have the power to keep the nose up below stall speed.
The first is commonly done by twisting the wing tips up at the trailing edge, reducing their angle of attack relative to the rest of the wing. It could also be achieved with a change of aerofoil section, or turbulators.
Both of these cause drag and reduce efficiency, and the second limits your ability to flare for landing. I imagine those were sufficient reasons for designers not to 'stall proof' their designs.

Bert Rutan's canard aircraft (Varieze, Long EZ, etc.) were designed to be spin proof. The  elevator is on the canard, in the front, and lifts the aircraft up, creating nose up moment, which balances the nose down moment from the lift on the wing. These aircraft are designed so that the canard, even with full aft stick, cannot create sufficient nose up moment to increase the AOA of the main wing to the point where the main wing would stall. The canard stalls first, loses lift, drops, begins flying again, and the cycle repeats. But the main wing never stalls or loses aileron effectiveness

@CharlesBretana And in what way is this different from a conventional design? The wing stalls first ... yada, yada, yada. Spin proofing is done exactly as described here, at least for moderate to high aspect ratios.

@CharlesBretana - that's basically the same as the elevator not having enough power to keep the nose up when flying at stall speed. You may want to ask a new question about why canard designs are not more common.

@Peter, the difference is that when a conventional wing stalls, you lose lateral roll control, and the tail (still producing nose up pitch moment) exacerbates the AOA and stall (even though the nose pitches down, the AOA is still increasing). In a canard, exactly the opposite happens.

In a canard, when the canard stalls, the wing is still generating lift, which produces a nose DOWN pitching moment, which reduces AOA on the canard, which breaks the stall on the canard.

There are other issues with a canard that cause it to not be as common, (e.g., Deep stall, when CG is too far aft and allows the canard to create high enough AOA to cause main wing stall)

From actual owners, landings are done crabbed into the wind and the aircraft corrects itself once both mains hit the runway.  It’s a lot like the airliners you see landing in a stiff crosswind.
Not really a dangerous design, but you have to adapt to its idiosyncrasies. Lots of dihedral makes it very stable and not being able to use the rudder uncoordinated will keep you from entering a spin.  Also, the elevator back-travel is limited to prevent the aircraft from stalling.  All of these built-in features can be overcome if you try hard enough, but flown within its parameters, it’s a safe ship.

You probably could spin it. A spin is just stall and yaw. If you pitch up, roll into 90 degree bank and pull hard into a stall, inertial coupling (as the flight path turns down away from the nose) will induce a yaw. the yaw will be upwards away from the earth, and so the lower wing will stall first, or more severely, causing a overbank that would induce the pilot to roll away. Unless the rudder is large enough to keep the yaw angle below the spin threshold, it just might spin...

@CharlesBretana It's enough to have asymmetric flow from turning. Spin will follow if stalled in a turn AND enough elevator authority AND a sufficiently backwards cg position.

@Peter, yes, Anything that creates yaw and high enough AOA to allow one or both wings to stall will by definition create spin conditions.


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Did Solar Impulse flight in commercial airspace and have kind of type 23 certification?
I read that the amazing Solar Impulse back in 2015 had circulated around the globe for 500 hours. It flew with speed 70 km/h and altitude of cruising 28000 ft. My questions are below

Did it fly in the commercial airspace? if yes then did it have kind of type 23 certification for small aircraft?
I have searched about the system of it such as flight control system and autoflight, but found nothing. What kind of the system was used in solar impulse? Did it have typical system with small aircraft such as Dassault Falcon or Pilatus?
Maybe anyone has detailed information related to solar impulse system such as flight control, autoflight, etc?

Thank you

I won't call an ATR a small aircraft. A bizjet such as a dassault falcon is smaller, not to mention GA.

You'd be amazed to learn how far you get with an experimental certification if you create enough publicity about that flight you try to accomplish.

@PeterKämpf so the experimental certificate allows the aircraft to fly in commercial busy airspace?

There isn't really such thing as commercial airspace.

@ZahiAzmi: Yes, all you need to do is to apply for a special permission and get that accepted. Publicity helps with that acceptance. A lot.

You are asking two completely different questions here: 1. about the certification and 2./3. about the flight control and autopilot system. You should have asked two independent questions for this.

I really cannot imagine it being anything else but "experimental aircraft"

You don’t need any type of special permission to fly a properly equipped experimental aircraft in airspace normally frequented by commercial traffic. In the US, the Classes of Airspace and even the non-military airways are open to any pilot and/or aircraft properly certificated (current license, medical, airworthiness, registration, etc.) regardless of if they are commercial, private, GA, etc. I fly out of four different airports in and under a very busy Class B. There are experimental aircraft that fly out of those same airports all the time. I fly right over the top of the B airport at TPA,

To answer the question presented in the title, I looked to Wikipedia for details on the flight(s) of Solar Impulse. According to Wikipedia, Solar Impulse 2 (HB-SIB) circumnavigated the Earth’s Northern Hemisphere in 17 legs/flights. Of those 17 legs, 14 legs had maximum altitudes above 18,000 feet MSL. In the U.S., Class A airspace is between 18,000 and 60,000 feet MSL. It is not considered “commercial” airspace. It is accessible to any aircraft with a Mode C transponder and ADS-b, including experimental aircraft. And, even this requirement can be waived with prior approval.

91.135 Operations in Class A airspace.
Except as provided in paragraph (d) of this section, each person operating an aircraft in Class A airspace must conduct that operation under instrument flight rules (IFR) and in compliance with the following:
(a) Clearance. Operations may be conducted only under an ATC clearance received prior to entering the airspace.
(b) Communications. Unless otherwise authorized by ATC, each aircraft operating in Class A airspace must be equipped with a two-way radio capable of communicating with ATC on a frequency assigned by ATC. Each pilot must maintain two-way radio communications with ATC while operating in Class A airspace.
(c) Equipment requirements. Unless otherwise authorized by ATC, no person may operate an aircraft within Class A airspace unless that aircraft is equipped with the applicable equipment specified in §91.215, and after January 1, 2020, §91.225.
(d) ATC authorizations. An operator may deviate from any provision of this section under the provisions of an ATC authorization issued by the ATC facility having jurisdiction of the airspace concerned. In the case of an inoperative transponder, ATC may immediately approve an operation within a Class A airspace area allowing flight to continue, if desired, to the airport of ultimate destination, including any intermediate stops, or to proceed to a place where suitable repairs can be made, or both. Requests for deviation from any provision of this section must be submitted in writing, at least 4 days before the proposed operation. ATC may authorize a deviation on a continuing basis or for an individual flight.

For further details about the aircraft, it’s custom autopilot and flight controls, and the actual flight Itself, visit https://aroundtheworld.solarimpulse.com/adventure .

For the solar impulse system:(question3)

The photovoltaic cell is made from two layers of Silicon
  (semiconductor material):
  - a layer doped with Boron which has fewer electrons than Silicon, this zone is therefore positively doped (zone P).
  - a layer doped with Phosphorus which has more electrons than Silicon, this zone is therefore doped negatively (zone N). When a photon(=
  smallest indivisible measure of energy associated with electromagnetic
  waves, ranging from radio waves to gamma rays) of light arrives, its
  energy creates a rupture between a silicon atom and an electron,
  modifying electrical charges. This is called the photovoltaic effect.
  The atoms, positively charged, then go to the P zone and the
  electrons, negatively charged, to the N zone. A difference in electric
  potential , that is to say an electric voltage, is thus created.
On the  Solar Impulse, there are exactly 17,248 silicon solar cells on
  a surface of 269.5m², capable of capturing up to 340kWh of solar
  energy per day. The energy created by these cells is stored in four
  high-performance lithium polymer batteries , located in nacelles
  isolated from the aircraft. The only drawback is the weight: they
  alone represent ¼ of the total weight of the aircraft.
These batteries power engines that have an average power over 24 hours
  of 15hp, comparable to a small motorcycle, and its maximum power is
  70hp. The four engines are fixed under the wings, with a two-blade
  propeller four meters in diameter. The total efficiency of this set of
  engines is 94%, which makes it a record of energy efficiency. The
  speed of the Solar Impulse varies between 36km / h and 140km / h,
  which is equivalent to the speed of a car

For the flight i think there is no rules ...see this:

us look at what happens during a 24 hour flight. For the solar panels,
  the day begins late and ends early: enough light to sustain flight is
  limited to just 10 hours per day! It’s a race against time...
6am: The sun has just risen. The airplane is on the runway. Its
  batteries, charged by the sun on the previous day, are nearly full. So
  it can take off using that stored energy... 6am-6pm: It gains
  altitude. The four motors turn at maximum power as the plane must
  create lift (the force that allows the plane to climb) to reach a
  thinner layer of air where there are fewer clouds. Despite the steady
  energy consumption, the batteries are charging. Nearing 6pm: It
  reaches 9000 meters (29,500 feet), its maximum altitude. The sun’s
  rays fade. The motors are throttled down and the plane starts to glide
  down to an altitude of 1500 meters (5000 feet), which takes about 4
  hours, during which time it consumes almost no electricity... Nearing
  10pm: At 1500 meters (5000 feet) of altitude, the pilot powers up the
  motors again, but this time they will take their energy from the
  batteries. The airplane flies like this until daybreak when solar
  energy again feeds power to the motors and recharges the batteries. A
  new cycle begins.
To determine the flight path that Solar Impulse will take in the sky, many factors must be considered: weather, air traffic areas, also the height of the land to be flown over and the performance of the airplane. A team of engineers and meteorologists based in Monaco Mission Control Center (MCC) determines the best route for the plane and then prepares the flight plan. Once the route is chosen, overflight and landing clearances must be negotiated for each country

The numbers in the picture are the energetic efficiency
http://solarimpulsemmktpe.e-monsite.com/pages/i-le-fonctionnement-et-la-conception.html
For more information :https://aroundtheworld.solarimpulse.com/adventure/technical-challenge-1

How does this answer the question about type 23 certification?

It is an awnser for the question 3

I am so sorry for not detailed question, by system, I mean "typical" aircraft system such as flight control or autoflight

Ah, sorry, I only read the title and the beginning of the question body. As stated, the question is way too broad though.

Question not clear

@L'aviateur sorry my bad

Yeah but is my question useless for you?

So i ad some info about the flight

See what's bold


text:
On the 737 MAX, can the pilots switch to the backup MCAS, or switch active MCAS input to another AoA probe during flight?
According to this article, my assumptions about the Boeing 737 MAX flight control system are (please correct me if I am wrong):

The aircraft has two flight control computers (FCC)
The aircraft has two angle of attack (AoA) probes
The Maneuvering Characteristics Augmentation System (MCAS) is implemented as a function of each FCC
Data from the port AoA probe is supplied only to FCC 1
Data from the starboard AoA probe is supplied only to FCC 2
Only one FCC (FCC 1 or FCC 2) is active during each flight, with the other serving as a backup system

Assuming a situation where only the AoA probe used by the active FCC produced incorrect AoA information and thereby causes the MCAS to intervene while the other AoA probe is functioning correctly, would it be possible to either

switch to the other FCC (which uses input data from the other AoA probe)

or

switch the active FCC to use input data from the other AoA probe

and thereby prevent the MCAS from continuing to use incorrect AoA data?

speculation about ongoing investigation are explicitly off topic

@ManuH It's no longer speculation. The final report for Lion Air has been available.

@JZYL you mean "it is no longer ongoing investigation". The question is still speculative, trying to know what would have happened if we rewrite history, event trying to know what would be reasonable to assume if we rewrite history. Less subjective formulation would be welcome.

"What if the system wasn't what it was but had the option (where?) to manually switch between FCC inputs to the Speed Trim, how would that system have performed?" Hard to know, since that depends on how well this band-aid to a band-aid got implemented. Totally opinion-based.

I changed the question to remove references and speculation about the two flights that crashed.
Note that my question was about whether or not the actual system design has the capability to let pilots switch between FCCs or ADRs, and not about how the system would have performed if its design were different.

why are there still downvotes

This question is now on about version 3, "would it be possible..." to which the answer is nearly always 'with enough time, effort, and money, yes.'  But why not just inhibit MCAS activation when the AoA inputs have significant differences between them for a significant time.

As I already mentioned, I considered my question to be about the existing design, not about possible changes to the design. Apparently, I did not phrase my question precisely enough. I apologize for the confusion.

@ABJX I suppose this is because people prefer to downvote rather than trying to point out what aspects should be improved.

we still have 2 downvotes. just to let the downvoters know.

As far as I understand from reading the B737 Operations Manual and the Lion Air Accident Investigation report,

Each Flight Control Computer (FCC) receives AOA input only from the onside ADIRU and vane.

The AOA and Mach inputs are provided to each FCC by the associated Air Data Inertial Reference Unit (ADIRU). Each ADIRU receives AOA information from one of the two resolvers contained within the associated AOA sensor (i.e. the Left ADIRU uses left AOA vane and the Right ADIRU uses the right AOA vane).
Source: FINAL KNKT.18.10.35.04 Aircraft Accident Investigation Report

The Maneuvering Characteristics Augmentation System (MCAS) and the Speed Trim System are controlled by the master FCC, which alternates between A and B per flight.

Similar to the Speed Trim Function, the MCAS function is also a flight control law contained within each of the two FCCs. MCAS is only active in the master FCC for that flight. At aircraft power-up, the master FCC defaults to the left side FCC; and will then alternate between the left and right FCC by flight.  The master FCC is not affected by the position of the Flight Director switches.
Source: FINAL KNKT.18.10.35.04 Aircraft Accident Investigation Report

This master FCC is different from the master FCC used to compute for the Flight Director (F/D) and the autopilot (A/P). For the purpose of AFCS, the master FCC can be cycled via the F/D switches on the Mode Control Panel (MCP), and the master FCC is illuminated.

The master FCC is indicated by illumination of the respective master (MA) F/D
  indicator light. The master FCC is determined as follows:

with neither A/P engaged in CMD, the FCC for the first F/D turned on is
  the master
with one or both A/Ps engaged in CMD, the FCC for the first A/P in CMD
  is the master FCC, regardless of which F/D is turned on first.

Source: Boeing 737-600/-700/-800/-900 Operations Manual

There does not appear to be a way for the crew to know which FCC is actively controlling the MCAS, nor a switch out the controlling computer, apart from pulling the circuit breakers.


text:
How does a helicopter phasing unit work?
I'm currently distance studying to become a B1.3 technician and the course material mentions a phasing unit in the control system, but the descriptions and images are very abstract and I'm having trouble understanding how it works. I've looked for other sources of information about the subject but have found next to none.
To my understanding the swashplate wants evenly distributed loads acting upon it, and the phasing unit somehow allows for this to happen. However, I don't understand what the phasing unit looks like, or its working principles, and to further confuse me; it can apparently be placed anywhere between the controls and the swashplate according to my coursebook.
Edit: I apologize to those who have responded; I should have uploaded an image from the start to be more clear.

I’ve worked with quite a few different helicopters and have never seen anything called a phasing unit.  Are you just asking about cyclic pitch?

Same here. Have written simulation physics models for about 20 helicopters, and have never heard of a phasing unit.

Is this the control mixing unit? https://aviation.stackexchange.com/questions/67322/why-do-some-helicopters-have-such-complicated-flight-control-mixers

There is a phenomenon called Phase Lag that is related to lead/lag hinge design, which has an effect similar to Precession Lag but varies at different angles less than 90 deg, whereas Precession Lag is always at 90 deg.  It sounds like the machine in question has Phase Lag issues that require active compensation in the control linkage, and there is a controller with actuators of some sort in the linkage.

@JohnK Yes there is phase lag in the blade flapping and lead/lag, usually fixed by placing the swash plate actuators at appropriate positions. So what would the phase unit actuators actuate? If it’s a mixer unit then yeah, we know what we’re talking about

@Jim No, I've got a handle on cyclic pitch and phase lag, and why it needs to be compensated. I've attached a picture of the text in my course book, which I realized I should've done from the start to avoid confusion.

As mentioned in the comments of the question, the function in question is normally performed by a control mixing unit (see this related question), also called control mixing or controller mixing. The principle is quite simple. Given the three actuators of your swashplate, you have to "map" the control inputs from the pilot's stick (so roll, pitch and collective inputs) to the three actuators given in your picture, such that the swashplate tilts or translates as requested.
Three examples:

The simples case is, when the pilot pulls on his collective, then all three actuators have to move down.

When the pilot gives a roll-right-input the desired output is that the swashplate tilts backward (seen in direction of flight). Therefore, in your diagramm actuator 1 has to move up, and actuator 2 and 3 have to move down a bit.

When the pilot gives a pitch-forward-input, the desired output is that the swashplate tilts to the left (seen in direction of flight). Therefore, in your diagramm actuator 1 stays constant, actuator 2 moves up a bit and actuator 2 moves down a bit.

Additionally such a mixing unit also automatically mixes inputs for example by automatically applying a rudder input whenever the collective is pulled.
It seems to me, that you textbook differentiates between these two actions. Mixing (e.g. automatically applying yaw when collective is pulled) is defined as mixing, and "phasing" is the act of distributing the mixed inputs to the servos spaced 60° apart.
As a reference look at this diagram of the control linkages of a Sikorsky S76, and observer that the mixing is performed in the same step as the "phasing".

Personally I have never heard of this, although I am involved in computerized helicopter control quite a bit. I would put such a function under "mixing". My opinion is that normally one does not differentiate between these two parts.

Your explanation makes a lot of sense. Thanks for your response. Your link just points to this page however, so I can't see the question you're referencing.
One thing I'm not sure about is the phasing unit and mixing unit being the same. The book talks about the mixing unit as well, as a completely separate thing.

@Smith: I corrected the link, but it is just the one from the previous comment. Concerning the mixing unit, could you take a photo of that page too? Perhaps we can work out what exactly the difference is.

Sure. Here are the next two pages in the book The mixing unit is only briefly mentioned, but there's a picture showing both mixing unit and phasing unit in a control system. Hope this clears it up.

Hm, it looks like I should update my answer. From what I can tell, the phasing unit (as the name indicates suboptimally) uses the previously mixed inputs and then allocates these to the three actuators. It seems to me that your textbook differentiates between mixing (e.g. yaw is applied automatically when collective is applied) and what they call phasing (distribution of the mixed inputs to the actuators...)

I am pretty sure the rotor always tries to align itself with the swashplate, so when the pilot gives a right-roll command, the swash-plate tilts right. Due to gyroscopic effect the blade pitch does indeed have to be raised forward (ccw rotation), but that's why the control horns ride the swash-plate ahead of the blades they control. If the rotor responded to swashplate tilting forward with tilting right, it would easily overshoot and exceed its flapping limits, because there is little or no torque back on the shaft. By matching the swashplate it won't tilt further until the body catches up.


text:
Did Helios Flight 522 take off with its pressurisation mode selector in manual?
I'm currently doing a write up on the Helios Flight 522 disaster. Full disclosure, this is a journalistic piece, and I'm not an aviation expert.
I'm many hours in at this point and most of it so far has made sense. The pilots missed the pressurisation mode selector being set to manual in pre-flight and take off checks, and then did not recognise and/or react appropriately the horn that went off in the cockpit. This resulted in rapid onset hypoxia, which incapacitated the pilots leading to the crash.
My problem is this. Alan Irwin, the ground engineer who was performed the pressurisation leak check before Flight 522 took off says that he in fact DID leave the switch in AUTO. In a recent Guardian article, he says:

"Except that it was not left in manual, says Irwin, who argues that investigators had not accounted for the distorting effects of impact."

and,

"Irwin’s explanation for the crash is that the system that controlled the outflow valve failed. But, of course, there was no proof either way."

So Alan is saying "hey, I left that switch in AUTO, and it was knocked to manual on the impact of the crash."
But, in page 66 of the Helios Flight 522 accident investigation report, it states that:

The NVM recording showed that on the accident flight, the cabin pressure control system was being operated in the manual mode.

So, according to the accident report, there is proof from the NVM recording that the switch was in manual BEFORE the flight crashed.
On a 737-31S, does the flight data recorder have a datapoint for the position of the pressurization mode selector? Cam we say with certainty that this plane DID NOT take off with the pressurisation mode selector in AUTO?
Why is Alan, a competent ground engineer who by all accounts seems to be very well-trusted, well-like and without incident throughout his career, saying that he thinks Helios Flight 522 crashed because 'the system that controlled the outflow valve failed'?
I also read the following in the accident report, page 102 (obviously this is referring to another earlier incident that took place on the accident plane):

On 16 December 2004, Helios Airways flight HCY 535 was en route from Warsaw, Poland, to Larnaca, Cyprus, cruising at FL350 at position EVENO, just prior to the top of descent to Larnaca, when the aircraft experienced a rapid decompression. The cabin altitude warning horn sounded and the cabin altitude went up rapidly. The passenger oxygen masks deployed automatically.

Okay, flight 522 didn't rapidly depressurise, it was never pressurised in the first place. But they go on to say:

The Air Accident and Incident Investigation Board (AAIIB) of Cyprus was not able to
reach a conclusive decision as to the causes of the incident, but indicated two
possibilities:
“a) An electrical malfunction caused the opening of the outflow valve; and
b) An inadvertent opening of the aft service door with the handle not in the closed
position.”

Is this what Alan is referring to? Can anyone help me make sense of this?

Welcome to Aviation.SE! I'm afraid your first question is opinion-based, which is off topic according to the help center. We cannot say why Alan Irwin said what he said.

Basically, what @Bianfable said. The only non-opinionated thing we can do is point you to the relevant parts of the official accident report - which you seem to already have.

Hmmm, that's fair. I guess I don't really know what I'm asking in a technical sense. I just wonder if it's a possibility that the pilots set off with the pressurisation mode selector in 'auto', but then switched it to manual when it became apparent that - as Alan claims - the system that controlled the outflow valve failed? The AAIIB states this very thing may have happened on the same plane. Or at least they think it may have happened.

I agree this is opinion based. He could be lying to try and avoid blame, or he could be right, or wrong in a different way. There's simply no way to give a definite answer. As for whether things are possible, thousands of hours went into the accident report, I can't see us adding technical insight.

My understanding is that it was actually NOT part of the SOP for the Ground Engineer to switch the pressurisation mode selector back to AUTO in any instance. You're right though, it could simply be him trying to rid himself of guilt. Still, I can't help but wonder if the NVM recording can show only that the pressurisation mode selector was being operated on manual mode at SOME POINT during the flight, but not whether or not the flight actually took with the pressurisation mode selector in MAN.

A non opinionated question might be "On a 73-[model in question] does the flight data recorder have a datapoint for the position of the pressurization mode selector".

Okay, no problem. I'll change it so it complies with rules.

The answer is on page 115 of the report. Sorry, I'm not sure how to delete this thread.

@NickBird No need to delete, you can post an answer to your own question.

The mis-positioning of a switch by ground crew should NOT have any bearing on an accident like this.  Before start, the Capt does a "flow" where all switches and controls are checked for proper positioning in a standardized tactile pattern, then the checklist is read through while visually confirming everything done during the flow.  There is no excuse for the flight crew not noticing that a switch was at the wrong position and even if the tech left a switch at MAN instead of AUTO, he is not to blame for anything.

I agree, and even the report agrees - it is absolutely the responsibility of the flight crew. Alan and his team were scape goats.


text:
Is it possible to upgrade old Mirage-III and Mirage-V variants with thrust vectoring engines?
Is it possible to upgrade old Mirage-III and Mirage-V variants with thrust-vectoring engines so that they can match the maneuverability of F-16s?

Adding thrust vectoring to arbitrary airplanes is pretty much like adding rear wheel steering to arbitrary motor vehicles. Yes you get super maneuverability but more often it's just unstable and dangerous.

Given enough tinkering, it is technically possible. But so many other modifications would be needed - strengthening, wing revisions to maintain the relationship between centers of lift and gravity, trim authority, managing the changes in stability, never mind the engine mounting and avionics refit, etc. etc. that really you would just be cannibalizing an old airframe to make an entirely new plane.
Somebody might conceivably do it as a short-life proof-of-concept prototype, if a new plane were proposed, the Mirage airframe was a good enough match, and a sophisticated digital control system was not necessary for these early tests. But in practice there is nobody in the world with all of the need, the industrial capability, the Mirages and the budget to embark on a new fighter programme that would require such a prototype.
Dassault did experiment with adding separate lift engines in the fuselage, in the Balzac V (which was just such a hacked-about Mirage III airframe) and the Mirage IIIV (which was twice the size), but it all went nowhere.


text:
Is it possible to use US-made engines in Russian made air-frames?
Suppose, someone buys 300 Pratt & Whitney engines from the USA and 150 Mig-35 air frames from Russia. Then, they use their in-house expertise to fit those engines to those air-frames.
What challenges would have to be faced by the user?

A really weird one, but... perhaps not off-topic enough to close?

It sounds like you are saying "if I have the expertise to do this thing, then can I do this thing?"  Naturally your expertise would include knowing how to cut any holes in the structure needed to make the engines physically fit, fairing over those holes (I'm thinking of the numerous "blisters" that appeared on the Bf-109), etc etc etc...

Rather than "legal" you might consider rephrasing to ask about the feasibility of getting them certified airworthy.  Assuming this is what you meant...

With enough time, effort, trial & error, and above all, MONEY, then all sorts of amazing & difficult (and even, ill-advised) things are possible.  Is it going to be easy? Nope.  Is it going to be cheap?  No, no, and absolutely NO!  Is it going to be worth doing? Highly doubtful!  But with enough billions of $, you could probably get it to work & get it certified as airworthy.  But really, what's the point of going to all that time, effort, and expense?  Just to say that you did it?

@RalphJ one reason I can think of is spares availability for the engines. Soviet engines are notoriously unreliable, needing far more maintenance than their western counterparts. This was a big reason for Germany and others to retire their almost new MiG-29s when joining NATO. If you can get the jets and engines on the black market, but no aircraft and correct engines for them, maybe it's your best option to make them fit.

Re-engine an existing airframe is such a complex undertaking, in a way that so much R&D is required that the entire aircraft ends up being reverse engineered to make it work. Or on some occasions, the airframe is first licensed or reverse engineered, after which, experience and expertise is gained to enable the re-engine effort.
China is notable in this area due their Russian lineage of airframes and a long lasting desire for western engines, and a mixed relationship with both. India is in a similar position but I'm not aware of any such attempts.
Notable examples:

Y-8-F600(Antonov An-12 clone) with PW engine, failed
Sabre II(Mig-21 derivative) with GE or PW engines), failed.
Shenyang J-11(license production Su-27 derivative) with Shenyang WS-10 engine (reverse engineered CFM56 derivative), successful.

We're not on law.stackexchange.com, but from what I can tell, the legal consequences would be limited to losing some benefits from warranty/support agreements.
There is no reason to do this. The engines on modern Russian fighters are quite good and among their selling points. You'd be concerned with upgrading their avionics - which are, of course, one of the most tightly-controlled export items - not their already excellent flying qualities.
This is not possible. For airliners or other pylon-engine configuration aircraft, re-engining is easy. For fighters with internal engines, it's not. The airframe is built around the engines, with special ring frames that allow room for the engine while carrying through the wing loads.
The engines are not geometrically compatible. None of P&W turbofans that could suffice to power a MiG-35 would fit in it.

TL;DR version: To paraphrase the old joke, you want a fighter that flies like a Flanker, sneaks like a Raptor, shoots like a Rafale, and maintains like a Gripen... not the other way around.
The MiG-35 is Russia's "low" component, similar to the F-16 in the US arsenal - lower cost, lower maintenance, high but not best achievable performance, and in its case also good improvised airfield capability. As part of that, it's a mid-size airframe, with smaller than average engines, optimized for efficiency and durability. The only US-built engine that could both fit and power the MiG-35 is the GE-414, after some modifications.
However, it's not the best possible engine for the MiG-35. The latest version of its own engine, the VK-10M, offers more thrust than the GE-414, has vectored thrust versions, and is a perfect fit for the frame. If you want the best, you'd want a custom engine, like an enlarged higher-bypass version of the EJ-200 (that's doable).
But, then, if you want the very best, you'd probably start with the Su-35 or the F-35. The MiG-35 is a workhorse.

If the question is not limited to specific airframes, the Irkut MC21 airliner is planned to have two engine options :  Pratt & Whitney PW1400G (American) and Aviadvigatel PD-14 (Russian).
So the answer is Yes, if you include airliners; you can even order them from the factory that way. First deliveries were scheduled for this year : I suppose that may be subject to change.
(there were also plans to give the Sukhoi Superjet a P&W engine option; these seem to have fallen by the wayside at the moment)

Designing an airframe to use multiple engine types had been done for decades (the 747 was launched with three different engine options from the major manufacturers) but that is a very different issue from retrofitting an engine that was never designed to be there. And it doesn't imply that you can buy the aircraft configured for engine X, and then change your mind and fit engine Y instead - not without some very extensive (and expensive) rework.

@alephzero all that is very true - and pod mounting makes the substitution vastly easier than in a fighter jet anyway. But unless someone has a line on 150 surplus MIGs (eBay?) I doubt the answers will have any practical relevance anyway.

You could then mention IL-96M (powered by PW2337). But this was still made by the OEM, so doesn't quite count in the spirit of the question.


text:
How can we compare the maneuverability of the F-16 Block 60 and J-10C?
Since J-10 originated from the IAI Lavi project, people assume that they are very close to F-16 in terms of maneuverability.

How can we compare the maneuverability of the F-16 Block 60 and J-10C?

A series of flight tests would be a good start.

@MichaelHall, is there any such test data available in the public domain?

Afraid not. Fighter manufacturers do release some performance data, but quite unsurprisingly it is limited, and not necessarily fully accurate.

@Jpe61, Chengdu J-10 vs F-16 E/F Fighting Falcon BLOCK 60

Yes, that site is based on data I described. The guys who put that  together did not test fly the planes. They have not even seen many (most) of them live.

OK, so it seems like you've got a comparison there. Is that unsatisfactory to you in some way or is that your answer? If it's an answer, put it down there in the "Answer" box and give yourself a check mark. If it's unsatisfactory, [edit] your post, include the link, and be explicit about what additional info you're looking for.

Block 60 isn't a good comparison because it's more of a ground attack airplane. The external thanks are heavy and breaks the aerodynamics.

@user3528438, which one is a better comparison?

I am curious about how the J-10 originated from IAI Lavi, because they are very different designs...

@Jpe61, U.S. Says Israel Gave Combat Jet Plans to China

Maybe block 50/52. See in your picture F16 Block 60 carries heavy tanks and an air-to-surface aiming module and no AA missiles, while the J10 carries all AA missiles and no external tank.

Thanks @user366312 this is new for me. Still I must say, they drifted very far from Lavi, very different configurations.


text:
If Sweden wants to export JAS39 to a 3rd country, will they have to get permission from the US govt.?
Does Sweden manufacture its own jet engine for JAS Gripen, or does it purchase engines from the USA?

JAS39 C/D uses Volvo RM12
JAS39 E/F uses Volvo RM16

I read on the internet that Gripen's engine is a modified American engine by Volvo. Does that mean that they purchase an engine from the USA and then modify it, or, did they purchase the design, and now they are independent for their own manufacturing?
I am curious to know if the USA's 2% export rules apply to these engines. I.e. if Sweden wants to export JAS39 to a 3rd country, they will have to get permission from the USA's govt.

"I read on the internet..." It might be useful if you include the source of this claim in your question

From your Wikipedia link: "General Electric produces 50% of the engine. Elements such as the fan/compressor discs and case, compressor spool, hubs, seals, and afterburner are manufactured in Sweden, final assembly also taking place there." What else do you want to know?

@Bianfable, I am curious to know if the USA's 2% export rules apply to these engines. I.e. if Sweden wants to export JAS39 to a 3rd country, they will have to get permission from the USA's govt.

Now I'm wandering if there is a single aircraft entirely built (enelctronic, hydraulics, tires, engine, materials, seats,...) by one country only... and its consequences on export restriction.

The most correct answer I can give is, It depends. More specifically, it depends on the specific components that are US sourced and it depends on who the ultimate end user is. But the most likely answer is yes.
All technology exports from the US fall under one of two sets of rules. They are the International Traffic in Arms Regulations (ITAR) controlled by the Dept of State and the Export Administration Regulations (EAR) controlled by the Dept of Commerce.
The US engines (F404/F414) that are the source of parts for the Volvo engines are military products controlled by the ITAR. When parts are exported separately from the complete engine, the export controls are set at the part level. So some parts may be controlled differently than others. But, in general any part controlled by the ITAR will require a license from the Dept of State before it can be exported. It is also possible that some components may be EAR controlled and require a license from the Dept of Commerce.
The licenses for export do contain restrictions on use, the end user, and reexport. That is typically spelled out in the individual license. Also, the 'technology' needed to transform the parts into an engine is controlled separately. That includes drawings, software, tooling, etc.
The question of reexport is covered in the ITAR under 22 CFR §123.9 Country of ultimate destination and approval of reexports or retransfers, and 22 CFR §123.10 Non-transfer and use assurances.
Specifically, from §123.9

(a) The country designated as the country of ultimate destination on
an application for an export license, or in an Electronic Export
Information filing where an exemption is claimed under this
subchapter, must be the country of ultimate end-use. The written
approval of the Directorate of Defense Trade Controls must be obtained
before reselling, transferring, reexporting, retransferring,
transshipping, or disposing of a defense article to any end-user,
end-use, or destination other than as stated on the export license, or
in the Electronic Export Information filing in cases where an
exemption is claimed under this subchapter, except in accordance with
the provisions of an exemption under this subchapter that explicitly
authorizes the resell, transfer, reexport, retransfer, transshipment,
or disposition of a defense article without such approval. Exporters
must determine the specific end-user, end-use, and destination prior
to submitting an application to the Directorate of Defense Trade
Controls or claiming an exemption under this subchapter.
(b) The exporter, U.S. or foreign, must inform the end-user and all
consignees that the defense articles being exported are subject to
U.S. export laws and regulations as follows:
(1) The exporter must incorporate the following information as an
integral part of the commercial invoice, whenever defense articles are
to be shipped (exported in tangible form), retransferred (in tangible
form), or reexported (in tangible form) pursuant to a license or other
approval under this subchapter:
(i) The country of ultimate destination;
(ii) The end-user;
(iii) The license or other approval number or exemption citation; and
(iv) The following statement: “These items are controlled by the U.S.
government and authorized for export only to the country of ultimate
destination for use by the ultimate consignee or end-user(s) herein
identified. They may not be resold, transferred, or otherwise disposed
of, to any other country or to any person other than the authorized
ultimate consignee or end-user(s), either in their original form or
after being incorporated into other items, without first obtaining
approval from the U.S. government or as otherwise authorized by U.S.
law and regulations.”


text:
What is the fundamental difference between a land-based jet and a carrier capable jet?
Are the carrier-capable planes designed as carrier-capable from the ground up? Or are they made carrier-capable by modifying a land-capable design?
For example, Mig-29 versus Mig-29K ?
In the latter case, what minimum changes, modifications, or upgrades are required to convert a normal jet plane into a carrier-capable plane?
For example, an F-16 block 50 is not a carrier-capable jet. Is it possible to make it carrier capable by applying minimum modification?

Specific modifications will depend on the country and type, but in general:
The obvious:

Adding a method of stopping the plane; typically a tailhook. While some land-based planes have a tailhook which can be used in emergency situations, sources seem to differ on the loads this is rated for. The tailhook on a naval aircraft is rated to stop it on a 300 foot long runway, land-based planes might catch the wire at the start of a runway and then take longer to decelerate for a lower load on the hook.
Space: Some aircraft are large, so features are added to make then smaller on the deck, such as folding wings or wingtips.
Adding a method to launch the plane. For some, catapults are used, typically requiring the nose gear to be able to withstand the load at launch. Other aircraft can be launched without catapults using ski-jumps, or simply just on a flat runway in full afterburner.
Depending on the aircraft, additional lift might be required, requiring larger flaps or bigger wings.

Strength:
The aircraft structures will be strengthened to account for much harder landings. According to one of the answers to this question, a normal carrier landing has a sink rate of approx. 600-800 feet per minute (FPM). This is in contrast to more normal landings where the goal is somewhere between 60-180 FPM.
This doesn't just impact the landing gear, the frame of the aircraft and every system on the aircraft must be capable of absorbing this added load on a regular basis.
Maintenance:
Carrier aircraft get exposed to salty water on a more regular basis than land-based planes. To avoid corrosion the aircraft have to be properly designed for, and treated to avoid it. I couldn't find sources with in-depth details, but I'd expect this to be a never ending task.
Aircraft carriers only have so much space, and sending an aircraft off to a depot for heavy maintenance if something breaks isn't really an option. Getting access to important parts of the aircraft is essential, so a land-based aircraft converted to fly from a carrier is likely to have more inspection panels facilitating access to key components.

See this related question:
Can an F-16 land on an aircraft carrier, at all?
Airframe, landing gear, tail hook, all much stronger. A carrier landing is mostly a controlled crash. A lot harder than a typical runway landing.
Air to air refueling must be probe and basket capable, vs boom only like on the F-16.
Of course the flight control software is different, but that's just software.

Are the carrier-capable planes designed as carrier-capable from the ground up? Or are they made carrier-capable by modifying a land-capable design?

AFAIK, designed from the start. I know of no jet that started as land-only, and then was 'modified' for carrier use.

@WPNSGuy  the de Havilland Sea Venom perhaps? https://en.wikipedia.org/wiki/De_Havilland_Sea_Venom

@Dave Gremlin - Good find! Probably the only candidate?

@WPNSGuy the F/A-18 started as the purely land based YF-17 and needed extensive redesign. The US Navy Fury was designed based on the land based F-86. The Soviet MiG-29K is a modified land based MiG-29, and they had versions of the Su-25 and Su-27 planned. A carrier capable F-16 version was planned and under design when the F/A-18 was chosen instead.

Refualing system is by convention only, there is no fundamental reason for a navy not to adopt the USAF system (or for an air force not to adopt the hose and drogue system, in fact it's in common use by air forces around the world).

@jwenting - Yes, "extensive" redesign on the YF-17 to turn it into the F-18.

@jwenting - for refueling, the larger aircraft NEED the boom, due to flow rate. Gassing up a B-52 via the drogue would take a LONG time. And the F-18 tanker is too small for a boom option. So, there mostly needs to be both.


text:
Did Turkey obtain external help in developing their jet engine?
Recently, Turkey unveiled a turbofan jet engine designated TF-6000. This engine is able to generate 6000 lbf of thrust, which is pretty good in my opinion, as the Taiwanese jet fighter, Ching Kuo, uses double 6100 lbf jet engines.
Did Turkey build this from its own research and know-how or did they obtain help from other countries, e.g., Ukraine?
Note: Ivchenko and Motor Sich engines have a much lower thrust than that of TF-6000

Why do you say that Ivchenko Progress' engines have a much lower thrust? E.g. the AI-222 has 5,500 lbf (>9,000 lbf with afterburner), which is similar. They also have engines with higher thrust (e.g. D-436 with 16,865 lbf), so why shouldn't they be able to help with something in between?

@Robe, I don't know. You tell me. If Ukraine had helped Turkey, they would have disclosed it to the media as their company's advertisements. The engines Turkey is purchasing have lower thrust. Other engines you are talking about do not match the amount of thrust.

Companies don't always disclose their cooperations, especially with sensitive issues like military capabilities. Perhaps Turkey wants to show the development as an inhouse achievement and asked Progress not to make publicity about it, but we can only speculate about it.

Your one-sentence question in the middle should be deleted and asked as its own question. We adhere to a reasonably strict "one question per question" policy - also, note that nobody has addressed it.

According to the article at:https://www.turdef.com/Article/tei-has-an-engine-surprise-for-kizilelma/1736

The engine will produce a trust similar to Ivchenko Progress and Motor
Sich’s AI-322F engine. The engine was planned to be used at Baykar
Technogy’s MIUS (Kızılelma) Combat Unmanned Aircraft System. The
engine is also suitable for the Turkish Aerospace ( TUSAŞ) Hürjet
Project. According to the TEI’s statement, the engine might also work
on supersonic platforms.
Baykar and Ivchenko-Progress had signed in November 2021 at
SahaExpo2021 AI-322F Turbofan Engine Supply and AI-25TLT Turbofan
Engine Integration contracts.

Thereby probably safe to say that some help from Ukraine in designing the engine was obtained.

Ivchenko and Motor Sich engines have a much lower thrust than that of TF-6000

@user366312 Currently the TF-6000 has no thrust at all - they are working on the prototype and the 6000 lbs are a paper value. OTOH, increasing thrust of an existing design a bit is very common in engine development.


text:
Is South Africa capable of designing and manufacturing turbofan engines for combat jets?
As far as I know, only a handful of nations and companies are capable of producing turbofan engines for combat jets:

GE, P&W, Honeywell (USA)
Safran (France)
Roles Royce (UK)
United Engine Corp (Russia)
Shenyang, Guizhou (China)
Motor-Sich, Ivchenko-Progress (Ukraine)
IHI (Japan)

The South African aerospace company Denel once designed and manufactured Atlas Cheetah jet fighters.
Did South Africa ever demonstrate the capability of designing and manufacturing turbofan engines for combat jets?
By "capability" I mean: If the above seven countries do not transfer technology, can South Africa build their own Turbofan engine? Does South Africa have enough human resources and money to invest to accomplish this?

Possible related: https://aviation.stackexchange.com/questions/49737/are-german-japanese-and-chinese-companies-not-able-to-build-aircraft-jet-engin/49827#49827

How do you define and determine "capability"? And why do you presume that they are NOT capable? (if they could make the Atlas Cheetah)  There is a difference between having design and manufacturing capability, and choosing to actually exercise it. (and to be commercially successful) For example: I am very capable of making a much tastier hamburger than McDonalds, (really, most people are...) but that doesn't mean that I would choose to start a business and compete for their customers. Or that I might make money at it.

you missed at the very least Volvo in Sweden who also build engines for jet fighters. Then there is HAL in India who build such engines under license from other companies. So they have the capability to build them at least, and can probably design them as well if they wanted to with some lead time.

SA managed to build and test a nuke. Turbine engines ........ a little more complicated and may have been completely manufactured for use in the Rooivalk attack helicopter but I must admit I was working on the armament targeting systems at the time, not primary power. Since SA was under some strong sanctions at the time I suspect that a few turbines were 'acquired' under the radar. Least said, soonest mended. officially the power source was a Turbomeca Makila

@jwenting "I'm sorry, Dave, I can't do that" is the only think that I could think of as I read your comment. :)

I think any answer is going to depend on politics.  What the apartheid-era South Africa could do, and perhaps more importantly, would do, is not necessarily what they could and would do today.

South Africa is quite capable of designing and manufacturing turbojet and turbofan engines from scratch. They have been doing so since the 1990's and the only reason they stopped was due to the upheaval after the end of the apartheid regime. They restarted the work in 2014, so who knows how far they've come since then.


text:
Is $664 million for 12 JF-17 block 3 reasonable?

Pakistani JF-17s replace Iraq’s F-16 fleet

iraq-might-buy-pakistans-jf-17-fighter-or-frances-rafale-or-both

... it’s rumored that Iraq has acquired 12 JF-17 Thunder Block III fighters, with the price tag speculated to be around $664 million.

This means each unit costs $55.33 million.
Isn't that expensive for a barely 4th gen jet?

iraq-to-reportedly-acquire-chinese-pakistani-jf-17-thunder-fighters

This report says that the value of the contract is $1.1 billion. This doesn't make any sense.

Presumably the $55 million unit cost includes not just the price of the jet (which is apparently around $25 million) but also armaments, spare parts, and support agreements?

Military jet purchasess, especially fighter jets, can't be boiled down to a simple formula like that.

When you change the oil in your car, you also change the filter, and replace a gasket or two.  When you read engine codes, you use a specialized computer to read them. The easiest way to do major work on a jet engine -- which they need periodically -- is to remove it from the aircraft.  Which requires specialized equipment, and then specific tools & replacement parts & testing gear to do that work.  And on & on.  Since the buyer wants more than just the 12 airframes (and when something breaks, the plane is done), the whole panoply of support equipment & spares inevitably has to come with.

define "reasonable"

@RalphJ: that sounds like an answer

When bidding for Inida's new fighter program in 2012, Rafale was only quoted about $100M each, which was considered very expensive already because F22 was just about that number though a decade earlier (and B2 was "only" about $1B each). The finalized price of Rafale was $2.44M eventually and that's just nuts-on-the-next-level. If you compare $50M JF-17 with PakAF's $40M J-10CE and yeah that's a hard sell, but if you see what's India and everyone else was paying for then I'd still call it a bargain.

Yes, it's reasonable.
International arms sales are usually done at significant markups - when they're not done for "free" (i.e. to attack the provider's opponent). Jet fighters aren't appliances, they are means of exerting control over some territory. These are things the purchasing country usually can't build on its own, at any cost.
The world's second-best-selling fighter, the Su-30 MKI, has cost India about 2x what it costs Russia. That cost is higher in large part because it includes assembly in India - and, unlike Ikea flatpacks and kit planes, military customers pay extra for a chance to build the product themselves. That markup is still low in comparison to 3x+ for India's Rafale deal.
Pakistan is on "enemy of my enemy" terms with the US, not too friendly with the EU or Russia, so China is its most reliable ally. The JF-17 is China's third-best fighter family, after the J-20 and Flanker clones. It's fitted with a full glass cockpit, a reasonable set of sensors, including IRST and reportedly a phased array in Block 3, can use most modern weapons, and has been successfully exported.
The F-16 - the best-selling fighter - has also been seen in deals with unusually high unit cost. Overall, it's comparable to the JF-17 in size, but better kinematically and with better avionics in Block 70/72, and costs $50M-$100M (too many sources to cite here). $55M for the JF-17, about twice its flyaway cost, seems in line with this.
Contract price is different from unit price. Arms sales contracts include maintenance equipment, spares, often weapons, and ancillary hardware. The $663M is for the planes and the rest of the $1.1B is for these other items.

I think block 3 actually uses AESA, otherwise they might just buy the block 2 instead.


text:
Why don't newer versions of F-16s use DSI inlets but newer Chinese jets prefer DSI?

In the above picture, we see that Chinese J-10 jets incorporated DSI inlets into their "B" and "C" versions.
(2) is the purported digital image of an Indonesian F-16 block 70/72.
However, F-16s retained MSI/MCID imtakes.
Why don't newer versions of F-16s use DSI inlets but newer Chinese jets prefer DSI?

while not adapted, it was tested: https://www.f-16.net/g3/f-16-photos/album11/album28/83-1120_001

Consider explaining what is DSI-- and likewise for the other abbreviations in the question--

J-10 has an adjustable intake which is essential for reaching its designed goal for speed. But it's heavy, complex, hence controversial from the early days (is the weight actually worth it?). Later when DSI is better studied and understood, it shows potential to be (almost) equally effective but a much lighter structure.
F-16 wasn't expected to be the best air superiority fighter to begin with, hence speed was neither a big requirement, nor a big problem (it has an excellent engine), so it chose a non-adjustable intake that is already very weight-efficient. DSI could potentially make it fly faster but that doesn't seem to have attracted a lot of interest (and investment).


text:
How does an altimeter deal with the non-linear pressure gradient?
Altimeters work via an aneroid barometer to sense the atmospheric pressure. If I climb from 2000ft to 4000ft, the altimeter senses that the pressure has reduced by, let's say, 70hpa, and is calibrated to know that 70hpa = 2000ft.
But the pressure gradient (or rate of change of pressure) reduces with altitude. A climb from FL240 to FL260 might result in a pressure reduction of only 30hpa. Yet it still knows that the aircraft climbed 2000ft. How does the altimeter do this?

As long as everyone’s altimeter uses the same 'pressure to reading' calibration, then, to some extent, the reading doesn't need to represent the actual difference in altitude. (See, for example, https://www.law.cornell.edu/cfr/text/14/29.1325, Part (e).)

Except that obstacle heights are not based on pressure altimeters. You still need altimeter accuracy when overflying an obstacle or high terrain.

Indeed: climbing from FL240 to FL260 represents a smaller pressure drop than climbing from, say,  2000' to 4000'. The instrument is calibrated for this (it is mechanically built in such a way to represent this pressure drop in a non-linear way) and can properly represent the measured pressure into a ISA calibrated altitude.
Just bear in mind that since altimeters have a fixed error margin in hPa, that hPA error will signify a greater error margin in feet as you climb. Which is a good reason why most GA altimeters are certified only up to 20.000 feet or so depending on how much you wanna spend on them. It is also the reason why there is an increased vertical separation (2000' instead of 1000') for flights above FL280.
eg. Regulations for TSO'd altimeters allow up to 20 feet of error at altitudes up to 1000', but allow for 140 feet or so at FL250

It uses a model called The International Standard Atmosphere (ISA) - so may be inaccurate if you don't have one of those to fly through :-)

license: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/3.0/

authors: Максим Копелян